JPH0632291A - Fly-by wire flight control system for aircraft and method for control of a plurality of positions of flight control wing face on aircraft - Google Patents

Abstract

PURPOSE: To select so that an operation of one flight control channel is sufficient to fly an aircraft by receiving a flight control wing face command and incorporating a means for connecting the command to a plurality of servo loops. CONSTITUTION: Before a flight control wing face command is executed, each ACE executes validity inspection for the command to confirm that a main flight computer relative to a flight control channel of the ACE correctly operates. If the inspection indicates that the main computer does not correctly operate, the ACE selects one set of flight control wing face command generated from another main flight computer, and uses it for controlling the set of the flight control wing face of the ACE. Accordingly, a left ACE 62 can control movement of the set 66 of the control wing face even when the computer 64 or a left data bus is faulted.

Description

Detailed Description of the Invention

[0001]

FIELD OF THE INVENTION This invention relates generally to aircraft control systems, and more particularly to redundant fly-by-wire control systems.

[0002]

BACKGROUND OF THE INVENTION Until the advent of fly-by-wire technology,
The flight control surfaces of commercial aircraft were controlled using complex system cables and mechanical controls as the main control path. Such a prior art control system is shown in FIG.
Partially shown in. In this type of control system, pilot control commands are transmitted from a pair of pilot controls 10 to a respective flight control surface 20 via a series of interconnected cables 12. The cable 12 drives one or more valves that control a plurality of hydraulic actuators 15, which actuators control the control surface 20.
To move. Cable 12 is pilot control 10
And a direct flight control wing surface 20. The failure auxiliary device 17 arranged at a plurality of important points allows safe operation of the system in the event of a cable failure.

The electronics bay controls several electrical or electrohydraulic actuators,
Provides improved control functions for the system. These actuators emphasize control commands entered by the pilot based on the flight conditions of the aircraft. Examples of such electro-hydraulic actuators include an outboard aileron lockout mechanism that prevents high speed aircraft aileron movement, an aileron droop mechanism that hangs the inboard aileron as a function of flap position, and the like. Other servo actuators include a series of autopilot servo actuators 14 that execute autopilot commands from an autopilot computer contained in the Electronics Bay.

The prior art control system shown in FIG. 1 has a number of drawbacks that limit its use in modern aircraft. The first drawback of such a system is its high maintenance cost. Electrohydraulic actuators, which often count forty or more on large aircraft, present a surprisingly large number of maintenance issues. Each of these devices is embedded within a complex path of cable 12 extending throughout the aircraft, so even simple repairs are labor intensive.

A second drawback of prior art control systems is the difficulty in implementing modern control methods that require a computer to control an aircraft. Since the introduction of these early flight control systems, advanced control methods have been developed to control speed, ascent and descent rates, bank angles, among others, while increasing aircraft stability. Such control methods are difficult to incorporate into mechanical control systems without substantially increasing system complexity. Finally, prior art control systems are inherently heavy. In aircraft design, it is always desirable to reduce airframe weight if done without reducing aircraft safety.
Thus, to overcome these and other limitations of prior art flight control systems, modern aircraft are designed to incorporate fly-by-wire technology.

In contrast to the mechanical flight control system shown in FIG. 1, a simplified diagram of a fly-by-wire (FBW) system according to the present invention is shown in FIG. In a fly-by-wire system, there is no direct mechanical connection between pilot control 10 and flight control surface 20. Instead of using a cable, the fly-by-wire system includes a set of pilot control transducers 22, which controls 10
The position of the pilot control 10 is detected and an electric signal proportional to the position of the pilot control 10 is generated. The electrical signals are transmitted to the electronics bay 24 where they combine with other aircraft data to generate flight control surface commands that control the movement of hydraulic actuators 26 that move the flight surface 20. 1 pair of pilot controls 10
Are normally connected by a failure aid 34 so that both controllers work together. However, if one of the pilot controls becomes stuck or fails, the other auxiliary pilot control is freed by providing the failure assist device 34 with sufficient force to uncouple the two controllers. Can be used.

Since safety is always a top priority in the aviation industry, fly-by-wire systems are typically redundant configurations that allow the pilot to still safely fly the aircraft if one component of the system fails. Contains elements. Such redundancy is provided for each axis. For example, one prior art fly-by-wire architecture has a separate system for controlling aircraft motion in each of the roll, pitch and yaw axes.

Each of the axis control systems typically included a main flight computer and a backup flight computer that controlled only the movement of the aircraft on a particular axis. If the main flight computer that controls the roll axis fails, the backup computer can engage to control the roll axis of the aircraft. Similarly, the pitch axis and yaw axis systems also include a main flight computer and a backup flight computer, respectively. However, if the axis channel backup computer fails, the other channel computers cannot function to fly the aircraft on that axis. Therefore, there is a need for an integrated fly-by-wire system that reduces the likelihood that an aircraft will not fly safely due to a failure of a portion of the system.

There is also a need for a fly-by-wire system in which each control channel in the system is divided into a series of independent control channels, each substantially separated from the other control channels. Thus, a malfunction that occurs on one channel does not affect the continuation of operation on the remaining channels.

Further, there is a need for a fly-by-wire system that includes multiple control channels designed such that the failure of some of the control channels does not affect the ability of the control channel to safely fly an aircraft. To do.

Finally, there is a need for a fly-by-wire control system that allows a pilot to fly an aircraft without the assistance of the flight control computer if all flight control computers included in the system fail.

[0012]

SUMMARY OF THE INVENTION The present invention includes a multiple redundant fly-by-wire control system for an aircraft. The system includes a set of pilot controls and a plurality of pilot control transducers connected to the set of pilot controls. Each pilot control transducer produces a position control signal proportional to the position of one of the pilot controls. A plurality of means for generating flight control surface commands receive position control signals and combine them with atmospheric data and data obtained from an inertial reference system to generate a set of flight control surface commands. To do. A plurality of actuator controller devices control movement of a set of flight control surfaces on the aircraft in response to the set of flight control surface commands received from the means for generating flight control surface commands. .
A set of flight control surfaces that are directly controlled by one of the actuator controller units will be sufficient to fly the aircraft if the remaining actuator controller units fail. Each of the actuator controller devices also responds to position control signals generated by the plurality of pilot control transducers in response to position control signals generated by the plurality of pilot control transducers in the event of a failure of the means for generating flight control surface commands. It is possible to control.
Included in each actuator controller device is a means for selecting a particular set of flight control surface commands to control the movement of the set of flight control surfaces. Each actuator controller device further includes a plurality of servo loop monitors for determining whether the plurality of servo loops controlling the movement of the set of flight control surfaces are operating properly.

[0013]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT A block diagram of the architecture of a fly-by-wire system according to the present invention is shown in FIG. The fly-by-wire system is divided into independent and separate flight control channels, which include a left flight control channel 60, a central flight control channel 80, and a right flight control channel 90. Since these control channels are completely and electrically isolated from each other, a failure in one channel does not adversely affect the operation of the other channel. As will be described in detail below, each of the flight control channels of the fly-by-wire system has a selected set of flight control wing surfaces of the aircraft such that the pilot can fly the aircraft using only one channel. To operate.

The fly-by-wire system includes a pilot controller 30 and a sub-pilot controller 32. Each of pilot controller 30 and sub-pilot controller 32 includes wheels 30a, 32a and control sticks 30b and 32b, respectively. Also included in the fly-by-wire system (not shown in FIG. 1) are other pilot controls such as a speed brake controller, a pair of pedals and a pair of lift and feel actuators. Pilot controller 30
Is connected to the auxiliary pilot controller 32 by the auxiliary device 34 at the time of failure. In normal operation, pilot controller 30 and subpilot controller 32 move together. However, if the pilot controller 30 or sub-pilot controller 32 fails, it is possible to free the other by providing sufficient force to the auxiliary device 34 in the event of a failure.

Coupled to pilot controller 30 and secondary pilot controller 32 is a bank of pilot control transducers 36 and a bank of secondary pilot control transducers 38. Pilot control transducer 36
And each of the transducers contained within the bank of sub-pilot control transducers 38
A plurality of pilot control transducer signals, which are proportional to the positions of the pilot controller 30 and the sub pilot controller 32, are generated. In the preferred embodiment, the bank of pilot control transducers 36 is the left set of pilot control transducers 36L (where the L after the two digit number in this specification should originally be lower case), pilot control transducers. Middle set of 36
c and the right set 36r of pilot control transducers. Similarly, the bank of sub-pilot control transducers 38 is preferably the left set of pilot control transducers 38L, the middle set of pilot control transducers 38c.
And a right set 38r of pilot control transducers. Each set of pilot control transducers produces a pilot control transducer signal, which is proportional to the position of the control stick or wheel to which the set of transducers are coupled. The pilot control transducer signals generated by the bank of pilot control transducers 36 and the bank of sub-pilot control transducers 38 are processed by actuator controller electronics (ACE) 62 on each channel to produce three separate data buses 40, 42 and Via 44 to two other independent flight control channels (60, 80 and 90). In the preferred embodiment, data buses 40, 42 and 44 are AR
INC629 digital communication links, which are standard in the aircraft industry, but other types of data communication links may also be used.

Flight control channels 60, 8
Since 0 and 90 are essentially identical, the following description of left flight control channel 60 applies equally to the center and right flight control channels. Therefore, the center and right flight control channels will not be described in detail.

The left flight control channel 60 is a left actuator controller electronic (ACE) device 6
2 and left main flight computer 64. Left AC
The E62 controls the movement of a plurality of hydraulic actuators that control the movement of a set of flight control surfaces 66, including symmetrically located spoiler pairs, ailerons, and flaperons, and the elevator, rudder and Stabilizer Flight control Controls the hydraulic actuator that controls the movement of the wing surface. Left ACE6, as described in more detail
2 sufficiently controls the control surfaces on the aircraft to allow the pilot to fly the aircraft if the central flight control channel 80 and / or the right flight control channel 90 fails. The other two flight control channels also have similar fault tolerant redundant control capabilities.

The left flight control channel 60 is powered by an independent power bus 68, which supplies power to the left ACE 62 and left main flight computer 64 via a set of leads 65. By using an independent power bus for each control channel, a failure on one power bus does not affect the operation of other flight control channels. The left flight control channel 60 also has an associated left independent hydraulic system 70, which is associated with line 72.
Used to power the movement of a set of hydraulic actuators (not shown). These actuators move the flight control surfaces included in the set of flight control surfaces 66. By using an independent hydraulic system for each control channel, failure of one hydraulic system is isolated so that it substantially affects only one flight control channel.

The pilot control transducer signal generated by the set of pilot control transducers 36L and 38L is transmitted to the left ACE 62 via lead 39L. Similarly, the pilot control transducer signal generated by the set of pilot control transducers 36c and 38c is transmitted to the central channel 80 via lead 39c, while the pilot control transducer signal generated by the set of pilot control transducers 36r and 38r is It is transmitted to the right channel 90 via the lead 39r. Left A to send and receive data from left data bus 40 in both directions
A set of leads 62L is used by CE 62, while a set of leads 62c and a set of leads 62r are connected to cause ACE to receive data only from central data bus 42 and right data bus 44, respectively.

After receiving the pilot control transducer signal from the set of pilot control transducers 36L and 38L, the left ACE 62 sends the signal to the left main flight computer 64 via the lead set 62L and the left data bus 40. The left main flight computer 64 receives pilot control transducer signals on a set of bidirectional leads 64L.
The left main flight computer receives pilot control transducer signals received from the left ACE 62 and pilot control transducer signals received from other control channels via the central data bus 42 and right data bus 44, atmospheric data and inertial reference system (see FIG. A series of flight control surface commands are generated based on data obtained from (not shown). The set of flight control surface commands generated by the left main flight computer 64 is transferred to the left ACE 62 via the left data bus 40.
Send to and return. The left ACE 62 then controls the movement of a set of hydraulic actuators (not shown) that respond to the set of flight control surface commands received by the flight control surface 66.
Can be moved.

The left main flight computer 64 receives the pilot control transducer signals from the central flight control channel 80 and the right flight control channel 90 so that even if one of the pilot control transducer sets becomes inoperative, the left main flight computer 64 will operate. 64 can still generate flight control surface commands. In addition, left ACE 62 receives a set of flight control surface commands generated by the main flight computer associated with central flight control channel 80 and right flight control channel 90. After receiving flight control surface commands from the main flight computer associated with other flight control channels, the ACE 62 uses which flight control surface commands to control the set of flight control surface surfaces associated with that control channel. Select whether or not.

As described below, each ACE typically selects a set of flight control surface commands generated by the main flight computer within the flight control channel of that ACE. However, before executing the flight control surface commands, each ACE performs a validity check on the commands to ensure that the main flight computer associated with the ACE's flight control channel is operating properly. If the validity check indicates that the main flight computer is not operating properly, the ACE selects one set of flight control surface commands generated by the other main flight computer to control the ACE flight control surface set. Used for. Thus, the left ACE 62 is still able to control the movement of the flight control surface set 66 in the event of a failure of the left main flight computer 64 or the left data bus 40. Above description is left flight control channel 6
0, but center flight control channel 80 and right flight control channel 9
0 is substantially the same in operation and failsafe function.

FIG. 4 is a functional block diagram showing how the fly-by-wire system according to the present invention interfaces with other avionics systems contained on an aircraft. As shown in FIG. 4, pilot controller 30 and sub-pilot controller 32 are coupled to a bank of pilot control transducers 36 and a bank of sub-pilot control transducers 38, which are proportional to the positions of pilot controller 30 and sub-pilot controller 32. Generate pilot control transducer signal.
Also shown is a set of pedals 33 used by the pilot and co-pilot to control the rudder (not shown) of the aircraft. Pedal 33 is also coupled to a bank of pilot control transducers 36 and a bank of secondary pilot control transducers 38. Pilot control transducer 36
Of the pilot control transducers generated by the bank of sub-pilot control transducers 38 and the bank of sub-pilot control transducers 38 through leads 39 to a plurality of actuator controller electronics (ACE).
Transmitted to. The pilot control transducer signals generated by the bank of pilot control transducers 36 and the bank of sub-pilot control transducers 38 are transmitted from ACE 62 to AC.
It is transmitted to the data bus associated with E.

Connected to the three data buses 40, 42 and 44 are three main flight computers,
These are a left main flight computer 64 associated with the left control channel 60, a central main flight computer 84 associated with the central control channel 80, and a right main flight computer 94 associated with the right flight control channel 90. As mentioned above,
Each main flight computer 64, 84, and 94 is A
A pilot control transducer signal is received from each of the CEs 62, 82 and 92. Each of the main flight computers is based on pilot control transducer signals and atmospheric data coupled to data buses 40, 42, and 44 via a set of leads 146 and data obtained from inertial reference device 145. Generate a set of flight control wing surface commands. The main flight computers 64, 84 and 94 then transmit the set of flight control wing surface commands via the data buses 40, 42 and 44 so that each of the ACEs 62, 82 and 92 receives a set of flight control commands from each of the main flight computers. The flight control wing surface command is received, and a plurality of power control devices 67
Select a particular set to control a, 67b and 67c. Each of the power control devices controls the movement of each flight control wing surface 20. As described above, a particular flight control surface controlled by one ACE, such as the left ACE 62, may be used if any of the flight control channels of the fly-by-wire system fails one or both of the remaining flight control channels. Is selected to allow safe control of aircraft flight.

Each ACE also receives via flap 101 a flap position count from a pair of flap logic blocks 100a and 100b. The flap position count corresponds to the physical position of the flap control wing surface of the aircraft and is generated by the bank of pilot control transducers 36 and the bank of subpilot control transducers 38 when the aircraft is flying without assistance of the main flight computer. ACE to adjust the gain of the pilot control transducer signal
The signals used by the are described below. .

Also shown is a direct mode switch 110 operable by a pilot or subpilot to disconnect ACEs 62, 82 and 92 from data buses 40, 42 and 44, respectively. If the direct mode switch 110 is activated, the pilot can control the aircraft in a direct analog mode, where the ACE controls the flight control surface movements without using the flight control surface commands generated by the main flight computer. is there. In this direct analog mode, the pilot must fly the aircraft without the use and benefit of advanced control methods implemented by the main flight computer. However, in the unlikely event that all main flight computers and atmospheric data and inertial reference devices 145 fail, the fly-by-wire control system of the present invention still allows the pilot to safely fly the aircraft.

FIG. 4 also illustrates data buses 40, 42 and 4.
4 illustrates other avionics components contained on the aircraft in communication with the vehicle. Aircraft Information Management System (AIMS) 120 receives signals from engine status display alert warning system 130 via a set of leads 131, both systems data buses 40, 42 and 44 via a set of leads 121. Be combined with. AIMS1
20 functions as a general purpose computer, which controls the functions of the aircraft such as flight management, manipulating navigation displays, displays indicating the need for onboard maintenance, aircraft communications, and collecting data regarding engine operation. Much of the data received and generated by AIMS 120 is data bus 40, 42 and 44.
To be shared with other flight system components by transmitting via. The set of direct leads 102 also provide AIMS 120 with flap position counts generated by flap logic blocks 100a and 100b.

The autopilot flight guidance system 140 provides computer control of the aircraft without the need for direct pilot or co-pilot input. With operation of the autopilot flight guidance system 140, the main flight computer generates flight control surface commands based on the signals received from the autopilot flight guidance system 140 instead of the pilot control transducer signals received from the ACE. Autopilot flight guidance system 14
0 also sends a backdrive signal to a pair of leads 142
Send to the set back drive servo motor. Back-drive servomotor is an autopilot flight guidance system 1
40, pilot controller 30, secondary pilot controller 32 and pedal 3
Move the set of 3 to correspond to the movement of the aircraft. This automatic movement of the controls provides a visual and tactile indication to pilots and co-pilots of how the autopilot system is operating an aircraft.

The pilot and sub-pilot controls 30, 32 and 33 are also banks of tactile devices 180.
Connected to. Roll and yawing feel device 180
Provides a fixed force for the displacement relationship such that the force required to move the wheel or pedal increases with displacement of the control. The variable pitch axis feel is generated by the pitch feel actuator 170. The pitch feel actuator receives signals from the ACE on a set of leads 171. The pitch feel actuator 170 modifies the force with respect to the displacement characteristics of the feel device 180 via a mechanical link 172. The pitch feel actuator 170 programs a pitch feel force that is proportional to the speed of the aircraft. Roll and yawing adjustment actuator block 19
0 is connected to the roll and yawing feel device via a mechanical link 191. The pilot and the sub-pilot receive the neutral position (the zero f) of the wheel or pedal via the adjustment command input to the adjustment actuator 190.
orce positon) can be changed.

A set of switches (not shown in FIG. 4) are
Lead 195 produces a stabilizer adjusted position signal that is transmitted to the ACE. The ACE then sends the stabilizer adjustment signal via data buses 40, 42 and 44 to the main flight computer. In response, the main flight computer generates a stabilizer wing command, which is sent back to the ACE. The ACE then sends this command on lead 192 to a stabilizer adjustment actuator (not shown), which controls the movement of the stabilizer flight control surface.

Finally, the standby attitude atmospheric data and inertial reference unit 145 also includes a set of leads 15.
Coupled to the data buses 40, 42 and 44 via 1 to provide failsafe redundancy in the event of a failure of the atmospheric data and inertial reference device 145.

FIG. 5 shows the internal components of the actuator controller electronics (ACE) 62. As discussed above, the actuator controller electronics receives the pilot control transducer signals generated by the set of pilot control transducers and executes the flight control surface commands generated by the main flight computer to generate a set of pilot control transducers. Flight control Controls the movement of the wing surface. Each of the actuator controller electronics (ACE) included in the fly-by-wire system according to the present invention is replaceable. A set of leads 202 for identifying the flight control channel of the fly-by-wire system in which the ACE is inserted.
The pin programming signal is provided above to the input signal management block 210 in the ACE 62. The ACE 62 has an internal power supply 220 connected to an independent power bus. A
CE 62 further includes a right data bus interface 230 connecting ACE 62 to right data bus 44, a central data bus interface 240 connecting to central data bus 42, and a left data bus interface 250 connecting to left data bus 40, respectively. Only one of the three data bus interfaces is ACE6
2 is bidirectional so that data can be sent and received from the data bus. The remaining data bus interfaces are "receive only" so that ACEs can receive but cannot send data to their data buses. ACE
62 further includes an internal data bus 260 through which the data in ACE is routed internally.

The pilot control transducer signals generated by the bank of pilot control transducers 36 and 38 shown in FIG. 3 are provided via lead 39 to block 270 which provides transducer support, signal selection, as will now be described.
And give flight control wing surface analog control. The transducer support and signal selection block 270 is also shown in FIG.
Flap logic blocks 100a and 100 shown in FIG.
The flap position count value is received from b through the lead 101. Transducer support and signal selection block 2
Included within 70 are circuits for exciting pilot control transducers and demodulating pilot control transducer signals received therefrom, and pilot control transducers or sub-pilot controls when the aircraft is flying directly in analog mode. A circuit for selecting a pilot control transducer signal generated by any of the transducers.

After receiving the pilot control transducer signal on lead 39, transducer support and signal selection block 270 transmits the pilot control transducer signal to multiplexer 280 via a set of leads 276, which is analog-digital on lead 281. (A / D) converter 290
Be combined with. A / D converter 290 converts the pilot control transducer signal from an analog format to a digital format and provides the digitized pilot control transducer signal to internal data bus 260 for transmission via left data bus 40 to the main flight computer. . The internal data bus 260 is the left data bus interface 250 and 2
It is connected to the left data bus 40 via the directional link 62L.

In the main flight computer (not shown in FIG. 5), the pilot control transducer signal is combined with the atmospheric data and data obtained from the inertial reference device 145 (not shown) using advanced control techniques, Generate a set of flight control surface commands used by the ACE 62 to control the set of flight control surface surfaces on the aircraft. The set of flight control surface commands is from the main flight computer to the ACE 62. Left data bus 4
0, left data bus interface 250 and bidirectional link 621. The flight control surface command set received by the ACE 62 is buffered in a signal command decode control block (SCDC) 300 before being converted from digital format to analog format in a digital-to-analog (D / A) converter 310. . As will be described further below, the SCDC block 300 uses a control signal transmitted on a set of leads 302 to control a plurality of sample and hold circuits 306. SCDC block 300 also controls the operation of multiplexer 280 using the control signals sent on lead 304.

A plurality of servo loops controlling the flight control surface are selectively connected to provide a sample and hold circuit 30.
A switch 320 is provided to receive either the analog flight control surface command output from 6 or the pilot control transducer signal generated by the bank of pilot control transducers from lead 276. The position of the switch 320 is shown in FIG.
Is controlled by a direct analog mode switch 110, shown in FIG. 1, which is coupled to switch 320 by an input signal management block 210 which controls lead 111 and switch 320 with signals on lead 212.

When the fly-by-wire system operates in direct analog mode, the switch 320 includes a plurality of servo loops 325, 330, 340, 350, 360.
And 370 are connected to connect to lead 276 instead of the output of sample and hold circuit 306.
In the direct analog mode, the set of flight control surfaces are directly controlled in response to pilot control transducer signals generated by the pilot control transducer, as will now be described. If the input signal management block 210 determines that the flight control surface commands generated by the main flight computer are invalid, or that there is a problem with all the main flight computers or data buses, then the input signal management block 21
A zero causes switch 320 to connect the input of the servo loop to lead 276 so that the fly-by-wire system operates directly in analog mode.

The ACE 62 shown in FIG. 5 for illustration in the left flight control channel of a fly-by-wire system provides flight control wing surface commands not only from the control channel's main flight computer (ie, the left main flight computer), but also in the center. It also receives from the central and right main flight computers using data bus interface 240 and right data bus interface 230, respectively. Left data bus 40 is left ACE
62 is the main source of flight control surface commands for 62. The central data bus 42 is the main source of flight control surface commands for the central ACE, and the right data bus 44 is the main source of flight control surface commands for the right ACE. If the input signal management block 210 detects a defect in the data received from the main data bus, the input signal management block 210 determines that the input signal management block is ACE.
Modifies the data bus that receives the flight control surface commands. If all main flight computers or all data buses fail, the input signal management block will
Switch 3 for CE to operate directly in analog mode
Change the position of 20.

Each ACE in the fly-by-wire system
Controls the movement of a set of flight control surfaces by providing flight control surface commands to a plurality of servo loops, each of the servo loops being a hydraulic pressure connected to one of the flight control surfaces. Control the actuator. Each ACE controls the elevator servo loop 325 that controls one hydraulic actuator on the elevator, the aileron servo loop 330 that controls the hydraulic actuator on the aileron, and the position of some of the spoiler wing surfaces on the aircraft. A set of spoiler servo loops 340, a rudder servo loop 350 for controlling a hydraulic actuator connected to the rudder of the aircraft, and a stabilizer adjustment control 360 for controlling the hydraulic actuator to move the stabilizer.
And an automatic speed brake arm or control actuator that controls the operation of the automatic speed brake actuator. Pitch speed sensor and monitor for providing pitch speed braking input to the elevator servo loop when the fly-by-wire system operates directly in analog mode 3
80 is also included in the ACE. Since the elevator flight control surface is one of the most important flight control surfaces on an aircraft, a method is provided by the present invention to ensure that the elevator operates correctly under all conditions.

FIG. 6 shows the input signal management (IS
M) block 210 and signal command decode control (SC
3 is a block diagram showing the operation of the DC) block 300. FIG.
Briefly, the input signal management block 210 determines whether the set of flight control surface commands received from each of the main flight computers (not shown) is valid and controls the servo loop associated with a particular ACE. Serves to determine which set of flight control surface commands to be used to select. The input signal management block 210 also determines whether the D / A converter 310, the A / D converter 290, and the internal data bus 260 are operating properly. If the input signal management block determines that there is a malfunction of all main flight computers, data buses, ACE D / A or A / D converters or internal data buses, the input signal management block 210 will allow the fly-by-wire system to directly analogize. Lead servo loop to switch 320 to operate in mode 27
Connect to 6.

The Signal Command Decode Control (SCDC) block 300 receives and stores the set of flight control surface commands generated by each of the main flight computers, and further identifies the selected set of flight control surface commands. Of A
Apply to the servo loop associated with CE. Which set of flight control surface commands is used is determined by the main flight computer (PFC) channel selection block 214 included in the ISM block 210.

When the main flight computer sends its set of flight control surface commands, it includes a series of cyclic redundancy check (CRC) words. CR
The C word is used by the main flight computer channel selection block 214 to determine if the data sent is valid. Typically, each ACE controls a flight control surface by selecting a set of flight control surface commands generated by the main flight computer in the flight control channel of that ACE. For example, left ACE
62 generally uses flight control wing surface commands generated by the left main flight computer 64. However, if the CRC word contained in the flight control surface command received from the left main flight computer indicates that the flight control surface command is invalid, then the PFC channel selection block 214 indicates that one of the other main flight computers has been selected. A set of flight control wing surface commands generated by

After selecting a set of flight control surface commands, PF
The C channel selection block 214 sends a signal to the SCDC block 300 indicating which set of flight control surface commands should be used by the ACE. The SCDC block 300 receives the set of flight control surface commands from all the main flight computers and stores them in storage register block 308. SCDC block 300
Address generator 301 contained therein receives a signal from PFC channel selection block 214 indicating which set of commands to use to control the servo loop associated with the ACE, and then outputs the signal via lead 305. The set of flight control wing surface commands that are transmitted to the storage register 308 are provided to a digital-to-analog (D / A) converter 310. The flight control wing surface command is converted from a digital format to an analog format by the D / A converter 310. The analog flight control surface command is then provided to the sample and hold circuit 306. Decoder circuit 309 also receives an address signal on lead 307 which causes the decoder circuit to activate the appropriate sample and hold circuit 306, so that the analog flight control wing commands output from D / A converter 310 are correct servo. Given to the loop.

As mentioned above, the position of the switch 320 may be selected by the direct mode switch 110 (not shown) by the pilot or subpilot, or by the input signal management block 210. If the PFC channel select block 214 determines that all primary flight computers (PFCS) are sending invalid data or no data at all, the switch 320 is toggled by the ISM block 210 so that the servo loop is Connected to receive the pilot control transducer signal directly on lead 276. In addition, the ISM block 210 has a wraparound monitor 217 that allows the D / A converter 3 to operate in the ACE.
10 or the A / D converter 290 is determined to be malfunctioning, the switch 320 is toggled.

Wraparound monitor 217 checks the operation of D / A converter 310 and A / D converter 290. When a flight control surface command is received by the ACE, the command is stored in a set of storage registers 216 included within ISM block 210. The set of flight control surface commands are converted from digital to analog in D / A converter 310 and sample and hold circuit 306.
Multiplexer 280 is selected to reconvert the flight control wing surface commands back to analog format from digital format in A / D converter 290 when applied to switch 320 via The reconverted flight control wing surface command is transmitted from the A / D converter 290 to the storage register 216 via the internal data bus 260. The wraparound monitor 217 then compares the reconverted digital flight control surface command to the original flight control surface command, which was also stored in storage register 216. If the two sets of flight control surface commands do not match within a predetermined error margin, wraparound monitor 217 declares an error in ACE and generates a signal on lead 212 to switch 32.
0 causes the servo loop to provide the pilot control transducer signal.

The multiplexer 280 also receives the servo loop status signal and other actuator information and
Used to convert to digital format by the / D converter 290 and send to main flight computer.

FIG. 7 illustrates a fly-by-wire system in accordance with the present invention that uses a redundant main flight computer, ACE and hydraulic actuators for rudder flight control surfaces 55.
8 is a flow chart showing by way of example how to safely control position 8 of FIG. In this example, the pilot actuates a control (rudder pedal) to produce a signal which is analyzed by the main flight computer and which is used to generate flight control surface commands to actuate the rudder.

The pedal 33 is coupled to a bank of pilot control transducers 36 which includes a set of separate pedal transducers P1, P2 and P3.
Each of the pedal transducers provides one of the ACEs with a pedal transducer signal proportional to the position of pedal 33. Then ACE 62, 82 and 92
One of each pedal transducer
The pedal transducer signal received from one of the two is transmitted to one of the data buses 40, 42 and 44. In particular, ACE 62 sends a pedal transducer signal from its associated pedal transducer P1 via left data bus 40, ACE 82 sends a pedal transducer signal from its associated pedal transducer P2 via central data bus 42, and ACE92. Transmits a pedal transducer signal from its associated pedal transducer P3 via right data bus 44. After the pedal transducer signals are applied to the data bus, the main flight computers 64, 84 and 94
Each of which selects a set of pedal transducer signals from the pedal transducer signals received on the data bus to generate flight control surface commands to move the rudder.

Again, since each of the flight control channels operates in substantially the same way, the following description will be directed only to the left flight control channel to simplify the discussion. After time t ₁ , when the ACE sends the pedal transducer signal onto the data bus, the left main flight computer 64 receives the pedal transducer signal from each of the three data buses 40, 42 and 44. At time t ₂ , the left main flight computer 64 selects one pedal transducer signal received from one of the three ACEs at block 375a for use in generating a set of flight control surface commands. To do. Central main flight computer 84 and right main flight computer 94 block 37 at time t _2.
The same operation is performed in each of 5b and 375c. Which particular pedal transducer signal is block 375
The choice of a is based on the selection of the intermediate value pedal transducer signal according to the voting rules well known in the aircraft control art. All other data received by the main flight computer on the data bus, such as atmospheric and inertial data, are selected in a similar fashion.

After selecting the pedal transducer signal from one of the three ACEs, at time t ₃ the left main flight computer 64 generates a set of proposed flight control surface commands at control strategy block 377a. . This is accomplished by combining selected pedal transducer signals with data obtained from the aircraft's atmospheric data and inertial reference system (not shown) according to a predefined control law for the aircraft. The actual control method used is derived from standard control method theory and empirical data collected during testing of the type of aircraft in which the fly-by-wire system according to the present invention is used.

After generating the proposed set of flight control surface commands at block 377a, the left main flight computer 6
4 sends the proposed command set via the data bus 40, while the central main flight computer 84 and the right main flight computer 94 center the proposed flight control surface command sets at time t ₄ . It transmits to each of the data bus 42 and the right data bus 44. After sending the proposed set of flight control surface commands to the data bus, the left main flight computer 64 sends the proposed set of flight control surface commands it generated to the other main flight computers 8.
Compare with the proposed flight control surface commands generated by each of 4 and 94. In the intermediate value selection block 379a, the left main flight computer 64 selects the middle value of each of the flight control surfaces command at time t _5. Block 3
79a, left main flight computer 64 at time t ₆
Transmits selected intermediate values of flight control surface commands via the left data bus 40, while the central and right main flight computers transmit selected intermediate flight control surface commands to the central data bus 42 and. Send on the right data bus 44. At time t ₇ , ACEs 62, 82 and 92 receive the set of flight control surface commands generated by each of main flight computers 64, 84 and 94. Time t
_{At 8} , the left ACE 62 is responsive to the set of flight control surface commands generated by one of the three main flight computers according to the signal selection function of the input signal management block 210 shown in FIG. 6 and described above. Select one of them. If the ACE 62 selects a set of flight control surface commands, it provides the selected set of flight control surface commands to the servo loop, which controls the rudder actuator 558a that moves the rudder on the aircraft. The operation of central main flight computer 84 and right main flight computer 94 is identical to that of the left main flight computer described above and need not be discussed.

FIG. 8 is an exemplary control loop diagram of servo loop control including power controllers controlled by respective ACEs for moving flight control surfaces 20. Servo loop control, generally indicated at 330, uses flight control surface commands to control a hydraulic actuator connected to one of the individual flight control surfaces 20, eg, an aircraft aileron. The flight control surface command is provided on lead 601 which is connected to summing block 602. Sum block 602 subtracts the actuator position signal on lead 627 from the flight control surface command on lead 601.
The actuator position signal is proportional to the position of hydraulic actuator 620 and indicates the position of flight control wing surface 20. The actuator position signal is generated by a linear variable differential transformer (LVDT) 622 that is connected to monitor the position of the hydraulic actuator 620. A position error signal representing the difference between the flight control surface command and the actuator position signal should be generated and transmitted on lead 603 to move the hydraulic actuator to the position governed by the flight control surface command. Indicates the distance.

The position error signal is provided via lead 603 to a second summation block 604, the output of which drives a servo amplifier 606. The output of servo amplifier 606 is provided on lead 607 to drive electrohydraulic valve 608. The electrohydraulic valve 608 controls the flow of hydraulic pressurized fluid to the hydraulic actuator 620, which moves the flight control wing surface 20. The LVDT position sensing transducer 610 is coupled to the electrohydraulic valve 608 and produces a valve position signal proportional to the position of the electrohydraulic valve 608. The output signal of LVDT 610 is coupled to demodulator 612 via lead 611. Demodulator 612 provides the demodulated valve position signal, which is provided via lead 613 to summing block 604 and back to complete the servo loop for electrohydraulic valve 608. The LVDT 622 also produces a signal that is transmitted on lead 623 to demodulator 626 and is proportional to the position of hydraulic actuator 620.
Demodulator 626 produces a demodulated actuator position signal and provides the signal to summation block 602 to complete the servo loop for hydraulic actuator 620.

Servo loop control 330 also includes a servo loop monitor block 636, which is an electrohydraulic valve 60.
8 and the operation of the two servo loops used to control the operation of the hydraulic actuator 620.
Servo valve monitor 614 reads from demodulator 612 to lead 613.
The demodulated valve position signal is received through and compared with the target valve position signal generated as a function of the position error signal provided by lead 603 to the servo valve monitor.
By comparing the valve position signal with the target valve position signal, the servo valve monitor 614 determines whether the electrohydraulic valve is accurately responding to the position error signal on the lead 603. If servo valve monitor 614 determines that electrohydraulic valve 608 is not operating properly, it produces an error signal, which is sent on lead 615 to servo loop monitor 636. The use of valve modeling to determine whether electrohydraulic valve 608 is operating correctly is as follows:
It is well known to those skilled in the aircraft control arts.

Upon receipt of the fault signal, the servo loop monitor 636 sends a valve open signal to the bypass valve 63 on the lead 617.
Send to 2. In response to the valve open signal, bypass valve 632 opens fluid valve 633 and fluid valve 634, thus blocking the flow of hydraulic fluid from electrohydraulic valve 608 to hydraulic actuator 620. When fluid valves 633 and 634 are open, hydraulic fluid that moves hydraulic actuator 620 can flow around closed path 644. This closed path allows the flight control surface 20 to be moved by another hydraulic actuator (not shown) as will be described. In addition to sending a signal to the bypass valve 632, the servo loop monitor 636 also sends a servo loop error signal to the multiplexer 280 shown in FIG. 6, which sends the servo loop error signal to the main flight computer to pilot the malfunction. Warning.

Demodulator monitor 630 checks the operation of demodulator 626 by comparing the output signal of demodulator 626 with the output signal of second demodulator 628. The second demodulator 628 is connected in parallel with the demodulator 626 and has an LV
The second demodulated output signal from DT 622 is provided to demodulator monitor 630. If the output signals from demodulators 626 and 628 do not match within a predetermined error limit, demodulator monitor 630 sends an error signal to servo loop monitor 636. In response to the error signal, the servo loop monitor 636 causes the bypass valve 632 to connect to the fluid valve 63.
Open 3 and 634.

The normal mode monitor 624 is also the lead 6
23 by monitoring the normal mode voltage on
Check the operation of DT622. If the normal mode voltage changes significantly from its specified normal range,
Normal mode monitor 624 sends the error signal to the servo loop monitor via lead 625. In response, the servo loop monitor causes bypass valve 632 to open fluid valves 633 and 634, thereby releasing the controls that PCU 330 has on flight control surface 20.

In addition to sending the servo loop error signal to the ACE's multiplexer 280 and opening the bypass valve 632. Identical to blocking valve 640 except that it is placed in another servo loop that controls other hydraulic actuators that are connected).

The blocking valve 640 is used for the flight control surface 20.
To prevent it from freely vibrating under certain failure conditions. Bypass valve 632 opens fluid valves 633 and 634,
The hydraulic actuator is no longer subjected to the force of the hydraulic pressurizing fluid from the electrohydraulic valve 608 and moves freely. If there are two independent hydraulic actuators connected to the flight control surface, the other actuator can continue to control the position of the flight control surface. However, if the other servo loop controlling the flight control surface also fails,
And if the bypass valve is opened, the flight control surface can freely swing. If the aircraft were flying at any high speed, the freely flapping flight control surfaces would combine with the larger flight control surfaces, resulting in catastrophic flutter. Therefore, the blocking valve 640 needs to block the position of the flight control surface. If a "blocking valve arm" signal is sent from another servo loop monitor to the blocking valve 640, and if the servo loop 330 fails, the bypass valve 632 in that servo loop will not open and will instead lock or block. Enter the mode. In the blocking mode, the closed path 644 is opened and the hydraulic actuator 620 is not free to move, thus blocking the flight control surface in place.

The operation of the servo loop monitor described above is independent of external loads placed on the flight control surface 20. An additional safety feature is added by testing the operation of the servo loop in the PCU without reference to the actual position of the flight control surface 20. Servo loop monitors must operate independently of the position of the flight control surfaces, as the very important flight control surfaces on the aircraft are driven by at least two hydraulic actuators, each controlled by a different control channel. . If the servo loop monitor does not work independently of the position of the flight control surface,
The fly-by-wire system may disengage the wrong hydraulic actuator if one of the actuators fails. For example, if a conventional command response monitor is used and one actuator fails, the flight control surface will not respond to flight control surface commands from the other healthy actuator. Therefore,
Command response monitors can deactivate healthy activators. This problem is eliminated by testing the elements that control the motion of the flight control surface regardless of the position of the flight control surface itself.

FIG. 9 shows a functional block diagram showing how the actuator controller electronics ACE62 directly controls the motion of the flight control wing surface (elevator selected for illustration) in analog mode. When operating in direct analog mode, the set of switches 320 is P
Open to prevent the CU servo loop and monitor 325 from receiving flight control surface commands generated by the main flight computer. Instead, the PCU servo loop and monitor 325 receives on lead 39L a pilot control transducer signal generated by a bank of pilot control transducers that is directly connected to a pilot control transducer (not shown). Each PCU servo loop substitutes the pilot control transducer signal for flight control wing command on lead 601.
Control the position of one of the flight control surfaces on the aircraft according to the control scheme shown in FIG. In addition, when operating in direct analog mode, the flap position counts from flap logic blocks 100a and 100b of FIG.
Lead 101 to variable gain block 275 contained within 0
Given in. Variable gain block 275 varies the gain of the pilot control transducer signal received from the set of pilot control transducers based on formulas well known to those skilled in the art of flight controller technology. PCU controlling aircraft elevator
For servo loops, the aircraft pitch velocity sensor and monitor block 380 further alters the gain of the position control signal before it is applied to the servo loop that controls the elevator flight control surfaces. Airplane pitch velocity sensor and monitor block 380 provides an aircraft pitch velocity braking signal that is superimposed on the pilot control transducer signal through lead 382 through summation block 274. The purpose of the pitch speed braking signal is to provide aircraft stability in the pitch axis. The output of summation block 274 is used to control the set of flight control surfaces when operating in direct analog mode.
Represents the sum of the pitch velocity braking signal and the pilot control transducer signal used by.

Switch 271 is responsive to the output signal of transducer monitor 279 to select a pilot control transducer signal from either the bank of pilot control transducers or the bank of sub-pilot control transducers. The transducer monitor 279 is the normal mode monitor 62.
4 and demodulator monitor 630 (illustrated in FIG. 8) to test the operation of these transducers.
In particular, the transducer monitor 279 has a lead 279a
A pair of demodulators 27 for receiving pilot control transducer signals on the bank of pilot control transducers on lead 279b and from the bank of sub-pilot control transducers on lead 279b.
7 and 278 are working correctly and ensure that the common mode voltage of the pilot control transducer remains relatively constant. The output of transducer monitor 279 is transmitted to lead 279c to control the position of switch 271.

FIG. 10 illustrates the distribution of flight control surfaces controlled by a particular ACE, and the connections between the set of transducers associated with either the pilot or subpilot controller for a typical aircraft. Generally, the controller transducers are distributed to the ACE so that each actuator receives input from its corresponding controller in direct analog mode. As mentioned above, the pilot's controller typically consists of wheels and control sticks, which are the auxiliary equipment 34 in the event of a failure.
To the wheel and control stick of the sub-pilot. In the preferred embodiment of the fly-by-wire system, ACE62 is further divided into Left-1ACE62A and Left-2ACE62B. The reason for this redundancy is
This is to increase safety and backup for the control of aircraft elevators and stabilizers as will be described. Coupled to the pilot's wheel and control stick are a set of wheel transducers 402 and a set of control stick transducers 4, respectively.
It is 06. Wheel transducer 402 includes three redundant transducers WL1, WL2, and WL3. Each of these wheel transducers produces a control transducer signal proportional to the rotational position of the pilot's wheel. Wheel transducer WL1 is left -1ACE6 by lead 404L1
2A, the transducer WL2 has a lead 40
4C is connected to central ACE 82 and transducer WL3 is connected to right ACE 92 by lead 404R.
Connected to. Similarly, the control stick transducer 406
Includes three transducers CL1, CL2, and CL3, each of which produces a control transducer signal proportional to the position of the pilot's control stick. Transducer CL1 is connected to left -1ACE 62A via lead 408L1, transducer CL2 is connected to central ACE 82 via lead 408C, and transducer CL3 is connected to right ACE 92 by lead 408R.

The set of pedal transducers 410 is
A control transducer signal proportional to pedal position is applied to each ACE. The set of pedal transducers further comprises three redundant transducers P1, P
2 and P3. Transducer P1 is connected to left -1ACE62A by lead 412L1 and transducer P2 is central AC by lead 412C.
E82, and transducer P3 is connected to right ACE 92 by lead 412R.

A set of speed brake transducers 41
4 is coupled to a speed brake lever controllable by both pilot and sub-pilot names. The set of speed brake transducers 414 includes four redundant transducers S1, S2, S3 and S4. Transducer S1 is left -1ACE by lead 416L1
62A, the transducer S2 has a lead 41
6L2 connected to Left-2 ACE62B, transducer S3 connected by lead 416C to central ACE8.
2 and the transducer S4 has a lead 41
Connected to right ACE 92 by 6R.

Connected to the sub-pilot control stick is a set of control stick transducers 418, each of which produces a control transducer signal proportional to the position of the sub-pilot control stick. The set of control stick transducers 418 includes three redundant transducers C
Includes R1, CR2, and CR3. Transducer CR1 is left -2ACE62B via lead 420L2
And the transducer CR2 is connected to the lead 420C.
Is connected to the central ACE 82 and the transducer CR3 is connected to the right ACE 92 by lead 420R.

Coupled to the sub-pilot wheel is a set of wheel transducers 422, which includes three redundant transducers WR1, WR2, and WR3. The transducer WR1 has a lead 42
4L2 connected to Left-2ACE 62B, transducer WR2 connected to lead 424C to center ACE
82, and transducer WR3 is connected to right ACE 92 by lead 424R. In addition to the pilot wheel and control stick controllers, both the pilot and co-pilot names are provided with stabilizer adjustment controllers for adjusting stabilizer flight control surfaces on the aircraft. Coupled to the stabilizer adjustment controller is a set of stabilizer adjustment transducers 430, which includes two redundant transducers LTA and LTC. Transducer LTA is left -1ACE6 by lead 432L1
2A is connected. Transducer LTC has lead 4
32C connected to central ACE 82. The co-pilot is also provided with a stabilizer adjustment controller that couples a set of stabilizer adjustment transducers 434 to it. Stabilizer adjustment transducer 434 includes two redundant transducers RTA and RTC. Transducer RTA is left-2ACE
62B and the transducer RTC is connected to right ACE 92 by lead 436R.

In the fly-by-wire system according to the invention, it is required that the transducers LTA and LTC, or transducers RTA and RTC, be in agreement before the stabilizer is moved. Transducer LTA produces a stabilizer adjust "arm" signal, while transducer LTC produces a stabilizer transducer "control" signal. Similarly, transducer RTA 434 produces a stabilizer adjust "arm" signal and transducer RTC produces a stabilizer adjust "control" signal. Therefore, in order to move the stabilizer, it is required that there is a match between the stabilizer adjustment "arm" and "control" signals. The ACEs of each of the three control channels, left-1ACE62A, center ACE82, and right ACE92, are provided with a position stabilizer signal proportional to the position of the stabilizer adjustment controller.

The set of stabilizer position transducers 438 includes three redundant transducers SP1 and SP.
2 and SP3. Transducer SP1 is connected to left-1ACE 62A by lead 440L1, transducer SP2 is connected to central ACE 82 by lead 440C, and transducer SP
3 is connected to right ACE 92 by lead 440R. The stabilizer position signals generated by the transducers SP1, SP2, and SP3 are transmitted by the respective ACEs on the data buses 40, 42, and 44 for indication to the pilot via the engine status display caution warning box 130 shown in FIG. Sent to.

Since the left ACE is split into two separate channels L1 and L2, the failure of the power supply to power ACE1 does not affect the operation of ACE L2 and adds safety to the control of the aircraft's elevator and stabilizer. . Listed in the columns of Tables 1 and 2 below are the specific flight control surfaces controlled by each ACE and the hydraulic actuators that move those flight control surfaces.

[0071]

[Table 1]

[0072]

[Table 2]

FIG. 11 illustrates the placement of flight control surfaces on a set of wings 500 on an aircraft. Main wings include one set of outboard ailerons 502 and 504, two sets of outboard spoilers 511-515 and 520-52.
4, one set of flaperons 530 and 532, and two sets of inboard spoilers 516-517, 518-
519. As can be seen from the above table, the majority of the flight control surfaces (except the spoiler) are positioned by two actuators, each controlled by a separate channel and separate hydraulic system of the fly-by-wire system. The letters contained within the circles adjacent to each actuator shown in FIG. 11 indicate the particular ACEs that control that actuator, while the letters contained within the square blocks indicate the actuator associated with each flight control surface. Shows a hydraulic system used by the. The main focus of this distribution of actuator hydraulic power and ACE control to move the flight control surface is a double hydraulic fault,
To ensure that no double ACE failure, or any combination of double hydraulic and ACE failures, reduces the maneuverability of the aircraft below safe levels.
This distribution also ensures that sufficient physical separation between the hydraulic system and the ACE control signal is maintained.

FIG. 12A shows the aircraft elevators and stabilizers and their associated ACE and hydraulic system arrangements. Each of the elevators 554 and 556 is controlled by two separate hydraulic actuators, which are controlled by different ACEs. The left ACE is further divided into two channels 62A and 62B so that a particular ACE controls only one actuator on both elevators 554 and 556. This division and increased redundancy will result in four ACEs.
It is ensured that the failure of any one of these cannot affect more than one of the actuators controlling elevators 554 and 556.

Aircraft stabilizer 552 is left -1AC
It is controlled by all four of E, Left-2ACE, Central ACE, and Right ACE, and uses the central and right hydraulics to provide hydraulic pressurized fluid. Aircraft stabilizer 552 is controlled using the stabilizer adjustment controls described above. To move the stabilizer wing surface, the stabilizer actuator is left-1ACE62.
A match is required between the flight control surface commands given to A and the central ACE 82, or between the right ACE 92 and the left-2ACE 62B. If these signal pairs generated by the ACE do not match, the stabilizer actuator cannot move the stabilizer blade surface. Matching is required between the two pairs of ACEs before the stabilizer moves, and current fly-by-wire systems require the left ACE to have two channels 62A and 62A.
Dividing into B gives this level of failsafe operation. By dividing ACE in this way,
Even if one ACE malfunctions, it is guaranteed that the stabilizer blade surface does not behave unexpectedly.
Stabilizer flight control surfaces must be properly controlled, as inadvertent movement, especially at high atmospheric velocities, can adversely affect the safe operation of the aircraft.

FIG. 12B shows the configuration of the ACE and hydraulic system used to control the rudder 558 of the aircraft. The rudder can be moved by any one of the three attached hydraulic actuators so that if any two hydraulic systems or any two
If one control channel fails, or if any combination of one hydraulic system and one control channel fails, it is possible to control the movement of the rudder 558. This fail-safe operation applies to any flight control wing on the aircraft, not just rudder flight control. Therefore, the system is capable of failing any two ACEs, or any two hydraulic channels, or any combination of one ACE and one hydraulic channel without losing the ability to safely fly the aircraft. Can withstand a breakdown.

Although the present invention has been disclosed with respect to a preferred embodiment, those skilled in the art of flight control systems will understand that modifications can be made to the system without departing from the spirit and scope of the invention. Let's do it. Therefore, the scope of the present invention is intended to be determined solely by reference to the appended claims.

[Brief description of drawings]

FIG. 1 is a schematic diagram of a prior art cable-based aircraft flight control system.

FIG. 2 is a simplified schematic diagram of a fly-by-wire control system according to the present invention.

FIG. 3 is a block diagram of a fly-by-wire control system according to the present invention.

FIG. 4 is a block diagram of a fly-by-wire control system according to the present invention.

FIG. 5 is a block diagram of an actuator controller electronics (ACE) included in a fly-by-wire control system according to the present invention.

FIG. 6 is an actuator controller electronic device (AC
FIG. 7E is a more detailed block diagram of E).

FIG. 7 is a flow chart illustrating the logic used by the fly-by-wire system according to the present invention to control the movement of individual flight control surfaces on an aircraft.

FIG. 8: Servo loops for individual flight control surfaces,
It is a figure with a servo loop monitor which determines whether the servo loop is operating correctly.

FIG. 9 is a functional block diagram showing how an (ACE) actuator controller device operates to control the motion of a flight control wing surface with the assistance of a main flight computer.

FIG. 10 is a diagram showing distributed signals from a plurality of pilot control transducers in an actuator controller device included in a fly-by-wire system.

FIG. 11 illustrates which actuator controller electronics and hydraulics included in a fly-by-wire system are used to control flight control surfaces located on a pair of aircraft wings.

FIG. 12 (A) shows which actuator controller electronics and hydraulics are used to control flight control surfaces on an aircraft elevator,
(B) is a diagram showing which actuator controller electronics and hydraulics are used to control the aircraft rudder.

[Explanation of symbols]

 36,38 Transducer 60 Left flight control channel 62 Left actuator controller electronics 64 Left main flight computer 66 Flight control wing surface 70 Left hydraulic system 80 Central flight control channel 90 Right flight control channel

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[Submission date] August 12, 1993

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Claims (6)

[Claims] 1. A fly-by-wire flight control system for an aircraft that monitors the position of a pilot control and generates flight control surface commands that control movement of a plurality of flight control surfaces on the aircraft, the system comprising: (A) includes a plurality of transducers associated with pilot controls, each of the plurality of transducers producing a signal indicative of the position of the corresponding pilot control; and (b) including a plurality of separate flight control channels, Each of the flight control channels includes: (i) a plurality of servo loops that control the movement of a set of flight control wings on the aircraft; and (ii) actuator controller electronics that receive signals from at least some transducers. Device (ACE), and (ii
i) a main flight computer coupled to the ACE and generating flight control wing commands as a function of the signal from the transducer, the ACE for each of the flight control channels receiving the flight control wing commands; And a means for coupling flight control surface commands to a plurality of servo loops, the set of flight control surfaces being controlled by each of the flight control channels, one set of flight control surfaces if the remaining flight control channels fail. A system in which the operation of the flight control channel is selected to be sufficient to fly the aircraft. 2. The fly-by-wire system of claim 1, further comprising a plurality of separate data buses, wherein all data transferred between flight control channels is transmitted on the data buses. 3. A flight control surface for monitoring the position of at least one set of replicated controls at a pilot position and a subpilot position and responsively controlling the movement of a plurality of flight control surfaces on the aircraft. Generate directives,
A fly-by-wire flight control system for an aircraft, comprising: (a) a plurality of transducers associated with a control, each transducer producing a signal indicative of the position of said control, and (b) a plurality of separated ones. An actuator controller electronic device (ACE) including a flight control channel, each of the flight control channels receiving (i) a signal indicating the position of the control.
And (ii) a main flight computer coupled to the ACE for generating flight control surface commands based on at least a portion of the signal, the ACE receiving the flight control surface commands and the flight control surface. Means for coupling the commands to a plurality of servo loops for controlling the movement of a set of flight control surfaces on the aircraft; and (c) the movement of the set of flight control surfaces is generated by the main flight computer. A system comprising means for bypassing the main flight computer and selectively providing signals directly to the servo loop for control without flight control surface commands. 4. A method for controlling the position of a plurality of flight control surfaces on an aircraft, the method comprising: generating a plurality of transducer signals at both pilot and subpilot positions, the transducer signals comprising: Indicating the state of control at each position, and selecting a different set of transducer signals from the plurality of transducer signals for each of the plurality of control channels; Transmitting to a separate main flight computer associated with each, and a plurality of corresponding flight control wing surface commands in combination with the set of transducer signals in each control channel with data obtained from the atmospheric data and inertial reference system. To generate a set of Steps and a set of flight control wing surface commands from the main flight computer are supported by a plurality of actuator / controller devices (A
CE), each of the ACEs selecting a set of flight control surface commands from a plurality of sets and providing the selected set of flight control surface commands to a plurality of servo loops. Controlling a set of corresponding flight control surfaces on the aircraft. 5. A step of monitoring each servo loop to determine if each servo loop is operating correctly and, if the servo loop is not operating properly, further opening a bypass valve in the servo loop. The method of claim 4, comprising. 6. The method of claim 4, further comprising the step of directly providing transducer signals to the plurality of servo loops in the event of a failure of a main flight computer that generates flight control surface commands.